The present invention relates to an actuator for an aircraft and, more particularly, to an actively controlled actuator for reducing vibratory transmissions from a gearbox mount to a support structure.
Helicopter main rotor lift and rotor driving torque produce reaction forces and moments on the helicopter main gearbox. The aircraft structure which supports the gearbox, e.g., transmission beams, are designed to react these loads and safely and efficiently transmit these primary flight loads to the airframe.
In addition to the primary flight loads, the aircraft is also subjected to vibratory loads originating from the main rotor system and acoustic loads generated by clashing oft he main transmission gears. These vibratory and acoustic loads produce vibrations and noise within the aircraft that cause discomfort to the passengers and crew. Low frequency rotor vibrations are a leading cause of maintenance problems in helicopters. Furthermore, as the aircraft reaches its maximum forward speed in level flight, the vibratory loads become very large, thus, producing increasingly high vibrations.
Many attempts have been made over the years to alleviate or reduce these vibratory loads and the resulting vibration and audible noise that develops within the aircraft cabin. A considerable amount of those attempts have been directed toward passive control of the vibrations. Some of the passive solutions have involved changes in rotor blade design to reduce the blade response to the periodic loading it experiences in forward flight. Other passive attempts have been directed toward reducing the transmission of vibratory and acoustic noise into the airframe or from the airframe into the cabin. For example, absorbing blankets have been incorporated between the airframe and the cabin interior for attenuating acoustic energy before it enters the cabin section. Another passive attempt involves the installation of low frequency vibration absorbers around the aircraft that are tuned to the vibration frequency of interest. The tuning is typically at a frequency of NP where N is the number of blades and P is the rotor rotational speed in cycles per second. Tuned absorbers have also been incorporated onto the main transmission support beams to produce a vibration impedance mismatch on and/or near the foot of the transmission.
One example of a passive vibration absorber is disclosed in U.S. Pat. No. 4,362,281 which relates to a pylon mounting system for supporting a helicopter gearbox. The pylon support is attached to the airframe substructure through resilient couplings or mounts. The couplings include elastomeric bushings which provide a soft resilient attachment between the pylon support and the airframe.
The above described passive solutions to reduce noise and vibration transmission have generally proven to be heavy and, consequently, not structurally efficient. These prior attempts also allow excessive motion of the gearbox causing gearbox-to-engine shaft misalignment under quasi-steady flight loads exerted by the main rotor.
There has been some recent attempts at producing active vibration and noise control systems. These systems monitor the status of the aircraft and/or the vibration producing component and attempt to command countermeasures to reduce the noise and vibrations. Active vibration and noise control systems are considered to be better at reducing aircraft vibrations and noise since the systems can be designed to counteract or cancel the vibratory and acoustic loads at or near the structural interface between the transmission and airframe, thus, preventing undesirable loads from entering the airframe.
U.S. Pat. Nos. 4,819,182 and 5,219,143 disclose one attempt at providing an active vibration control system. This system includes a plurality of vibration sensors, e.g., accelerometers, that are located at strategic places throughout the aircraft and provide signals to an adaptive control unit. The control unit provides signals to electro-hydraulic actuators that are located within a series of struts which support the gearbox. The actuators produce controlled forces which attempt to minimize vibration at the sensed locations.
Another active control system is discussed in U.S. Pat. No. 5,588,800. This active control system is mounted within a helicopter rotor blade and includes actuatable flaps on the rotor that are controlled to reduce the blade vortex interaction and/or vibratory loads transmitted to the airframe.
Many of the active control systems that are currently being evaluated or proposed utilize hydraulically operated actuators to provide the counteracting forces for damping the sensed vibratory loads. These actuators include a piston arrangement that is attached to a mounting stub through a ball or universal joint. These types of joints, however, tend to bind under high load, especially high vibratory loads. Also, these actuators incorporate conventional internal seals which are not suitable for vibration and noise. As such, the seals quickly wear out and are not very efficient at attenuating acoustic noise.
Recently, a vibration reduction system was incorporated into an EH-101 aircraft manufactured by Westland Helicopters. The system included a semi-stiff strut mounted in parallel with the piston load to carry the large quasi-steady flight loads. In this design, however, the strut provided a path for undesirable high frequency acoustic vibrations.
A need therefore exists for an improved actuator for use in an active vibration control system to minimize NP vibratory and high frequency acoustic transmissions from a vibrating component into the aircraft airframe.
The present invention relates to an active vibration and noise control system for controlling the transmission of vibratory loads from a vibrating component. The control system includes an actuator that is designed to attach the vibrating component, such as a gearbox mounting foot, to a support structure. The actuator is actuated by the control system to control the transmission of steady-state and transient loads from the vibrating component.
In one embodiment of the invention, the actuator includes a housing and a mounting member. The housing mounts to the aircraft. The mounting member attaches to the vibrating component within the aircraft. A lap-fit piston arrangement is attached to the housing and includes a sleeve located within the housing, and a piston slidably disposed within the sleeve.
An elastomeric bearing assembly engages the mounting member with the piston. The bearing assembly includes a first bearing located between the mounting member and an inner bearing member. The first bearing is adapted to transmit axial loads between the mounting member and the piston. A second bearing is located between the inner bearing member and the housing and is adapted to transmit moment and shear loads from the mounting member to the housing. A third bearing is located between the inner bearing member and the piston, and is adapted to permit rotational movement of the inner bearing member relative to the piston.
A diaphragm separates the piston from the bearing assembly so as to inhibit hydraulic fluid from contacting the bearing assembly.
The control system includes a processor which receives a plurality of signals representing the vibratory state of the component or airframe and the position state of the actuators. The processor sends signals to a hydraulic actuation system for controlling the movement of the actuators.
The foregoing and other features and advantages of the present invention will become more apparent in light of the following detailed description of the preferred embodiments thereof, as illustrated in the accompanying figures. As will be realized, the invention is capable of modifications in various respects, all without departing from the invention. Accordingly, the drawings and the description are to be regarded as illustrative in nature, and not as restrictive.